The present invention relates to a device for connecting an external electrical power supply source to an aircraft on the ground.
It is standard practice, in particular to avoid an excessive noise level in airports, to replace the internal electrical power supply system of an aircraft on the ground by an external source, such as a static power supply or a power supply using one or more electrical generating sets. It can indeed be understood that, if the aircraft auxiliary power units (or APU) have to be shut down on the ground, it is necessary to continue providing an electrical power supply without discontinuity to some of the aircraft equipment, including the lighting and heating systems, and especially the in-board computers which must not suffer any electrical power supply interruption, even a brief one.
In the past, aircraft were equipped with battery banks whose purpose was to continue supplying at least the in-board computers during the time interval elapsing between the two successive operations which were then performed consisting first in disconnecting the in-board 400 Hertz auxiliary power units, and then in connecting the external source to the electrical installations of the aircraft, this connection being in fact achieved by connecting the external source to a connecting socket provided for this purpose underneath the aircraft.
In order to save weight and volume, these back-up batteries have been eliminated on aircraft of recent model and replaced by an electronic measuring, computing and control device whose purpose is to perform connection of the external source and correlative disconnection of the in-board generators with a very small overlap time under phase agreement conditions which should, in principle, enable any break in the power supply to the in-board computers to be avoided, whilst not giving rise to any damaging phenomena.
FIG. 1 is a schematic block diagram of an installation of this type, such as is fitted at present to a certain number of aircraft of recent model.
The part which is comprised inside the aircraft and which is surrounded by a dashed outline is designated in this figure by the reference 1.
The aircraft conventionally includes at least one in-board auxiliary power unit 2 which delivers a 400 Hertz alternating current on a three-phase electrical power supply line 3 supplying the in-board equipment.
Supply of the line 3 is achieved via a three-phase switch X2, whose opening and closing are controlled by a measuring, computing and control unit 4. This unit 4 is connected by a control input 5 to the aircraft cockpit. Control of the switch X2 is achieved via a connecting line 6.
The central power supply line 3 is also connected, via another three-phase switch X1 similar to the switch X2, to an external multipin socket 7, generally located underneath the aircraft, which is designed for connection to an external source. The switch X1 is controlled by the unit 4 via a connecting line 8.
Inside the aircraft, the external socket 7 is in addition connected to the measuring, computing and control unit 4 by a connection 9. The computer and control unit 4 measures the external voltage applied to the multipin socket 7. The unit 4 includes a two-wire output 10 which supplies power to two of the multiple pins of the external socket 7.
The external source 25 is, for example, connected to a three-phase distribution bar 11 and supplies 400 Hertz current in the airport. The distribution bar can comprise several feeders Dn-1, Dn, Dn+1, . . . , the feeder, line or cable, Dn for example being assigned to power supply to the aircraft 1 oh the ground.
The cable Dn is connected to a multipin connector 12, which is a connector fitted to the aircraft socket 7,via a three-phase switch 13, normally electromechanical, whose two-wire control connection 14 is connected to two of the terminals of the multi-pin connector 12. This connection is achieved in such a way that, when the connector 12 is fitted to the conjugate socket 7, there is a continuity of electrical connection between the two-wire connection 10 and the two-wire connection 14. Opening and closing of the switch 13 are thus controlled by the unit 4 via its output 10.
In flight and during a certain time after landing, the line 3 is supplied with electrical power by the on-board APU 2, the switch X1 being open and switch X2 closed.
When the aircraft on the ground is to be supplied from the external source via the cable Dn, the connector 12 is connected to the external socket 7 of the aircraft The switch 13 is then open, so that this connection is made with the connector 12 de-energized.
From the cockpit, a power supply transfer order is then sent, via the connecting line 5, to the control unit 4, which then first supplies the two-wire connection 10 so that the two-wire connection 14 is also supplied, and closes the three-phase switch 13. The connector 12 and socket 7 are then supplied with electrical power by the external cable Dn
The unit 4 then detects, via the connection 9, the presence of a three-phase voltage on the socket 7. It consequently controls the APU 2, via the connection symbolically represented by 15, to phase them with the external voltage, taken on the three-phase measuring line 9. In present day aircraft, this phasing is considered to be satisfactory by the unit 4 when it is achieved to within 90 degrees
The unit 4 then orders closing of the switch X1 followed, approximately 60 milliseconds later, by opening of the switch X2. The line 3 is then supplied by the external source only, after a 60 ms overlap with the internal source 2.
This way of proceding has nevertheless proved dangerous, in particular (but not only) when the external source 25 is provided from the electrical mains by means of static converters.
Indeed, parallel connection of two three-phase sources 25 and 2, which has a duration of about 60 ms whereas these two sources are only in phase to within 90 degrees, can result in electrical power exchanges between them which can have the following consequences :
automatic shutdown of the external source 25 on overload, PA1 destruction of the external source 25 if electrical power is supplied to it by the internal source (APU) 2, the static converters not being able to absorb power, PA1 breaking of a shaft in the APU 2 due to stresses gene by the suddent increase of the power to be delivered by the APU, which is even more serious as it affects the aircraft itself.
Subsequent replacement of the external source by the internal source 2 is achieved by the reverse operations phasing of the APU 2 with the external voltage to within 90 degrees, closing of the switch X2 and opening of the switch X1 60 ms later, then opening of the external switch 13 and disconnection of the connector 12. During the 60 millisecond overlap period, the damaging phenomena described above can occur in exactly the same way. The object of the invention is to overcome these drawbacks.